1. Field of the Invention
The present invention concerns a method for aiming a satellite with respect to a celestial or heavenly the sun or a star for example, and a device suitable for implementing it.
Satellite in this case means any artificial object moving in the solar system. This object may in particular be
in an orbit around the earth or any other planet in the solar system, PA1 in an orbit around a satellite of any planet in the solar system, or PA1 in a solar orbit, possibly a transfer orbit between two planets. PA1 ensuring the recharging of the batteries with which the satellite is equipped, PA1 powering the various items of equipment required by the satellite: sensor, computer, heaters, remote control and telemetry, in particular; and PA1 ensuring illumination of the satellite by the sun so that the thermal configuration of the satellite is homogeneous and maintains the equipment within the permitted temperature range. PA1 seeking the sun by rotating the satellite about one axis (for example the roll axis), so that the field of view of at least one solar sensor encounters the sun; PA1 rotating the satellite about one axis (for example the pitch axis) so as to bring the direction of the sun towards the desired direction; PA1 rotating the satellite about the direction of the sun, and controlling this direction so that it becomes identical with the desired direction (for example the roll axis). PA1 U=demand signal to be applied as determined by the control law PA1 S.sub.s/c =unit vector of the instantaneous direction of the heavenly body in question (sun or star) calculated from the angular measurements of the sensors PA1 SR=unit vector of the reference direction of the heavenly body in question PA1 .omega.=measured velocity vector PA1 C=velocity of rotation demanded about the direction SR PA1 Kd=velocity regulation gain PA1 Kp=position regulation gain PA1 .LAMBDA.=vector product (or, in English, cross product). PA1 U=demand to be applied PA1 SR=unit vector of the reference direction of the sun PA1 SR.sup.T =transposed unit vector of the reference direction of the sun PA1 .omega.=measured velocity vector PA1 C=rotation velocity demanded PA1 Kd=velocity regulation gain PA1 Kp=position regulation gain PA1 e.sub.M =measurement axis vector PA1 N.sub.sy =solar sensor measurement PA1 N.sub.by =correction of the solar sensor measurement PA1 I=identity matrix PA1 L=limitation factor PA1 .LAMBDA.=vector product (or, in English, cross product). PA1 complexity of the control laws, which are very different from the conventional laws of the earth and sun acquisition modes; PA1 poor accuracy of aiming about the non-measurable component axis since this axis is controlled passively by coupling and the error in attitude about this axis is never measured or determined; and PA1 difficulty in using this control law to generate thruster commands on the aforesaid two axes since that requires: PA1 U=demand to be applied PA1 SR=unit vector of the reference direction of the sun PA1 .omega.=velocity vector measured PA1 C=velocity of rotation about SR PA1 Kd=velocity regulation gain PA1 the necessity for two additional stages: pitch rotation to bring the sun facing the sensor used, and then yaw rotation to cancel out the yaw error due to the particular choice of the reference direction of the sun; and PA1 poor accuracy of aiming, which deteriorates over time because of drift in the gyrometers and because the control law does not include the positional control term. This poor aiming accuracy: PA1 which is applicable to any triple-axis stabilized satellite whatever the arrangement of its thrusters (unlike French Patent No. 2,649,809), PA1 which does not call into question the logics conventionally used for the sun, earth or star acquisition modes (unlike French Patent No. 2,659,809), PA1 which uses the measurement of a single solar or stellar sensor with a single measurement axis (unlike French Patent No. 2,407,860, European Patent No. 0,338,687 and U.S. Pat. No. 5,080,307) but in combination with other available measurements; PA1 which can easily be programmed in the on-board computer; and PA1 the aiming accuracy of which is optimum because of a positional control carried out actively and continuously, by comparing the calculated direction of the heavenly body with the direction aimed at or reference direction. PA1 the axis of sight belongs to a sensor with at least one sensing axis, the first reference plane being defined as being perpendicular to the sensing axis; PA1 the second reference plane is perpendicular to the first reference plane and is defined as containing the axis of sight and sensing axis; PA1 the sensor is a single-axis sensor; PA1 the sensor is a twin-axis sensor, a single output of which is used; PA1 the first and second quantities representing the first and second angles are the tangents of these angles; PA1 the control law is of the type: EQU U=-Kd*[.omega.-C*S.sub.s/c ]-Kp*[S.sub.s/c .LAMBDA.SR] PA1 U=demand to be applied to the torque generator PA1 S.sub.s/c =unit vector of the instantaneous direction of the celestial object PA1 SR=unit vector of the aiming axis forming the rotation reference axis PA1 .omega.=measured velocity vector PA1 C=rotation velocity demanded about SR PA1 Kd=velocity regulation gain PA1 Kp=position regulation gain PA1 .LAMBDA.=vector product (or, in English, cross product) PA1 the celestial object is the sun; PA1 the aiming axis is an axis which is at least approximately close to an inertia axis of the satellite, chosen so as to obtain continuous illumination of a solar generator installed on the satellite, by virtue of which the satellite is in a sun-aimed mode; PA1 the satellite has another sensor with a second axis of sight, suitable for detecting another predetermined celestial object, and the aiming axis is chosen so as to form, with this second axis of sight, an angle at least approximately equal to the (sun) - satellite - (or other celestial object) angle; PA1 the other celestial object is a star, by virtue of which the satellite is in star acquisition mode; PA1 the other celestial object is the earth, by virtue of which the satellite is in earth acquisition mode; PA1 the celestial object is a star; and PA1 the satellite also has a terrestrial sensor with a second axis of sight, and the aiming axis is chosen so as to form, with the second axis of sight, an angle at least approximately equal to the star-satellite-earth angle, by virtue of which the satellite is in the mode for the acquisition of the earth from a star. PA1 the sensor is a sensor with a single sensing axis, the first reference plane being perpendicular to the sensing axis, and the second reference plane containing the sensing axis; PA1 the sensor is a solar sensor; PA1 the satellite also has a stellar sensor; PA1 the satellite also has a terrestrial sensor; PA1 the sensor is a stellar sensor; PA1 the actuating unit includes thrusters. PA1 the unit for measuring the instantaneous rotation velocity of the satellite includes gyrometers.
2. Description of the Prior Art
As is known, satellites have, for the purpose of controlling their orbit and attitude, sensors and actuators connected to logic processor units within a system normally referred to as the attitude and orbit control system. The logic processor units are often integrated within an on-board computer, the actuators including for example thrusters (and/or magnetic coils, and/or gears) while the energy consumed on board is provided by batteries, solar cells subjected to solar radiation and/or propellants used for thrust propulsion in particular for orbit control.
As for the sensors, these are in practice of several types, depending on the nature, size, brightness, etc of the heavenly body (earth, sun or star) which they serve to detect. There are currently single-axis or dual-axis sensors.
Single-axis sensor means in general an appliance providing an angular coordinate of the direction of the heavenly body aimed at in a frame of reference peculiar to the sensor. In practice an axis of sight and a sensing axis are defined for the sensor, and the angular coordinate is for example the angle made by the axis of sight of the sensor to the projection of the direction of the heavenly body aimed at onto the plane containing the axis of sight perpendicular to the sensing axis.
A dual-axis sensor means an appliance supplying two angular coordinates of the direction of the heavenly body aimed at in a frame of reference related to the sensor (which determines this direction completely). The appliance therefore has the function of two single-axis sensors with identical axes of sight and separate sensing axes, generally orthogonal. For detecting the sun, such a dual-axis sensor is in fact normally formed by associating two such single-axis sensors.
In fact a satellite may be caused to adopt several attitudes during its life after being released from its launch vehicle.
Thus, for example, the nominal attitude of a satellite of the triple-axis stabilized type moving on a circular terrestrial orbit consists of having a Z axis, known as the yaw axis, pointing towards the earth, a Y axis known as the pitch axis, perpendicular to the plane of the orbit, and an X axis known as the roll axis, perpendicular to the Z and Y axes and having the same direction as the instantaneous linear velocity of the satellite on its orbit, the direction of the Y axis being such that the frame of reference (X, Y, Z) is positive. Such a nominal attitude is controlled by means of a terrestrial sensor, either alone or combined with solar sensors or a stellar sensor.
Other types of attitude may be envisaged, notably just before a satellite is put into its operational orbit, or after a serious failure of any type affecting in particular the attitude and orbit control system.
It is normal in such cases to attempt to put the satellite into a so-called sun-aimed attitude in which it is in slow rotation about an axis pointing towards the sun, chosen so as to be close to a principal inertia axis and such that the satellite solar panels are illuminated. This enables the satellite to await subsequent commands, while ensuring its safety, that is to say:
Conventionally, this attitude control mode is referred to as the sun acquisition mode and is based on a sequencing of the following type given by way of example:
The satellite thus being in a sun-aimed attitude, the requirements of the mission generally require it to come (or return) to its nominal attitude, that is in practice for it then to be aimed (or re-aimed) towards the earth (or a star). The method of seeking the earth or star which is conventionally used, in such an earth acquisition mode or star acquisition mode, consists, during the seeking stage, of putting the satellite into slow rotation about an axis aimed towards the sun, this axis being chosen so that the field of view of the earth (or star) sensor necessarily encounters the earth (or star).
In some cases, when it is not possible to seek the earth directly, a reference star is initially sought, the Pole Star for example: once the star has been found as described above, the satellite is, from the measurements of the star sensor, controlled with respect to rotation so as to bring the field of view of the earth sensor facing the earth: this is then referred to as acquisition of the earth via the star.
The rotation axis pointed towards the sun during the seeking of the earth or star forms, with the axis of sight of the earth or star sensor, an angle at least approximately equal to the sun-satellite-earth angle or sun-satellite-star angle. This axis may be the direction which the sun should have in the frame of reference related to the satellite once the latter is in its attitude pointed towards the earth or star.
These different attitude control modes therefore require the direction of the sun or star in question to be determined for the purpose of taking it into account in the control loops using the control law corresponding to the current stage of the attitude control mode: in particular in the last phase of the sun acquisition mode, in the first phase of the earth or star acquisition modes and in the last phase of the acquisition of the earth via a star.
Prior art Patents which relate to such a change in nominal attitude, after an intermediate change into a sun-aimed attitude, are French Patent Nos. 2,407,860 and 2,649,809 (MESSERSCHMITT-BOLKOW-BLOHM), European Patent No. 0,338,687 (BRITISH AEROSPACE), and U.S. Pat. No. 5,080,307 (HUGHES AIRCRAFT).
In the first three patents, the direction of the sun is determined from the measurement of two single-axis solar sensors whose sensing axes are perpendicular, or from a twin-axis solar sensor.
Likewise, in European Patent 0,338,687, the direction of the star is computed from the measurement of two single-axis stellar sensors whose sensing axes are perpendicular, or from a twin-axis stellar sensor.
The computation is carried out by determining the three coordinates of the unit vector of the instantaneous direction of the sun or star in the frame of reference of the sensor from two angular measurements defining the orientation of this direction in the frame of reference.
Control is effected by applying demands of the type: EQU U=-Kd*[.omega.-C*S.sub.s/c ]-Kp*[S.sub.s/c .LAMBDA.SR]
where:
During the sun, earth or star acquisition modes, the velocity of the satellite about its three axes is measured (by means of gyrometers, for example) and used in the attitude control laws, solely for the purpose of damping the positional control or carrying out the velocity control.
An earth-seeking mode using a single solar sensor is described in French Patent No. 2,649,809. This patent proposes a method in which the direction of the sun is not completely determined. The control law is based on the fact that the rotation of the satellite causes coupling between the errors in attitude according to the measurement axis and according to the non-measurement axis. Thus, when the single-axis solar sensor detects an error, the control law generates a command suitable for cancelling out this error according to the measurement axis, while the non-measurable error on the other axis is eliminated by coupling.
The control law used is of the type: EQU U=-Kd*[.omega.-C*SR]+Kp*[SR.LAMBDA.e.sub.M +(SR*SR.sup.T)e.sub.M -e.sub.M ]*L(N.sub.sy -N.sub.by)
where:
It will be noted that this command includes a term relating to velocity control according to three axes and a term relating to positional control according to two axes, namely the measurement axis and an axis perpendicular to this measurement axis and to the reference direction of the sun, referred to as the non-measurable component axis.
This method has the following drawbacks:
either a complicated logic for generating impulses of different durations on the different thrusters: the impulse modulator conventionally used on known satellites is then not applicable, PA2 or using thrusters having specific orientations suitable for producing torques on both axes, which de-optimizes the system of attitude control of the satellite by thrusters since the directions of these thrusters are then imposed by the datum of these two axes. PA2 prevents its application to the sun acquisition mode, the long duration of which (typically several hours) would, with this type of law, result in latent aiming errors, PA2 prevents its application to the star acquisition method, which requires great aiming accuracy because of the need to recognize the star; and PA2 the risk, in some cases, of resulting in failures of the earth acquisition.
An earth-seeking mode using a single solar sensor based on coupling between axes due to the rotation is also described in the publication: "The attitude and orbit control of the EUTELSAT II spacecraft" .sctn.6.11 page 95; Symposium on Automatic Control in Space - IFAC - 17-21/7/1989. This document proposes an earth-seeking mode in which the reference direction of the rotation vector of the satellite during earth seeking is chosen so that it forms, to the direction of the sun, an angle equal to the earth-satellite-sun angle and so that its component which cannot be measured is zero. This article deals with the pitch component, which amounts to saying that the sun is maintained in the XZ plane of the satellite throughout the earth-seeking phase and that the final attitude at the time of sensing the earth has a yaw angular difference. During earth seeking, only velocity control about the direction SR is suggested, so that the control law used by this method is probably of the type: EQU U=-Kd*[.omega.-C*SR]
where:
This method has the following drawbacks:
On the other hand, the present invention relates to a method for aiming the satellite towards a heavenly body such as the sun or a star: